Rotor speed control for rotorfixed wing aircraft



June 8, 1954 K. HOHENEMSER 2,680,579 ROTOR SPEED CONTROL FOR ROTOR-FIXED WING AIRCRAF Filed March 26, 1951 3y Cam/L GM 7%? MM flnam/s s.

Patented June 8, 1954 UNITED STATES PATENT OFFICE some .srErn QONTROL FOB. RQTQB- rixnp WING AIRCRAFT Kurt H. Hohene nser, Rock Hill, Mo., assignor to McDonnell Aircraft Corporation, St. Louis, Mo., a corporation of Maryland Application March 26, 1951, Serial No. 217,572 4 Claims. (01. 24417.13)

This invention relates to rotor speed control varied in response to the relative speed of the means for substantially autorotating rotors for rotor shaft and the forward speed of the aircraft, aircraft, and is of particular value in aircraft of including means for permitting the pilot to con the type utilizing a Substantially autorotatin trol the rotor. lifting rotor along with a small "fixed wing for With the above and other objects in view, this fiightconditions of large rotor advance ratio when invention resides in the novel features of form,

all or most of the driving torque for the rotor construction, arrangement, and combination of is provided by the relative airflow over the airparts presently described and pointed out in the craft. claims.

In Such xed win ai craft, Whi e e li In the om an in .drawin is gradually built up and the rotor lift reduced Fig. 1 is a graph showingthe eiiect on rotor with increasing flight speed, the rotor speed must torque of changing the collective blade pitch nevertheless be kept within certain limits. A angle as compared with tilting the effective plane sudden loss of rotor speed at a high flight speed of the rotor,

resulting from increased flight speed, maneuver- Fig. 2 is a schematic plan view of a rotor-fixed ing, or encountering gusts, is likely to cause a wing aircraft of the type referred to herein; and

failure of the rotor blades and a loss of the air- Fig. 3 is a schematic elevational layout of a craft. Hence, there must be no uncontrollable de rotor speed control system embodying and exceleration of the substantially autorotating rotor. emplifying the present invention.

It has been found that the rotor speed in a Referring now by reference characters to the rotor-fixed wing aircraft in the conditions of drawings, which illustrate a preferred embodiautorotation or close to autorotation is highly ment of the invention, in Fig. '1 the coefiicient of sensitive to changes in forward flight speed, and rotor torque Co is plotted against the blade pitch the speed of rotation is likely to decrease dangerangle 13 for various angles A of the swashplate ously with increased forward speed. This undewith the relative wind. A constant advance ratio sirable characteristic is one of the ain reasons of 0.2 is assumed.

why relatively little has been done in recent years In Fig. 1, positive values of C indicate a to advance the development of this type of airdecelerating rotor torque, while negative CQ craft.

values indicate an accelerating rOtor torque. From The principa o ject of the r ent invention th fi ur it is ap arent hat n sp o. is to provide automatic speed control for sub? celerating or decelerating the rotor, the tilting of stentially ne ft e r tor wh ch Wil t s la chang n e an le A) has a efiective under high rotor advance ratios and powerful and reliable eifect. With regard to under all such gust and maneuvering conditions change of rotor blade angle B, it is to be noted as are ordinarily encountered. that the effect of such change is neither power- Another object of the invention is to provide ful nor consistent. Between pitch angles of zero means, to be utilized at low forward speeds, for degrees and about three degrees, an increase in eliminating the automatic speed response of the angle accelerates the rotor, but from approxirotor. mately three degrees upward, a further increase The invention consists in the provision of 40 in pitch angle decelerates the rotor.

means for governing the speed of an autorotating In the high speed flight range of such a rotorrotor by adjusting the longitudinal tilt of the effixed Wing aircraft, gusts and maneuvers will fective plane of the rotor with respect to the relacause chan s n angl o atta k of th Order of tive wind responsive to changes in rotor speed. ma nitude of ten degrees. Referring to Fig. 1, it

The invention also consists in the provision of i apparent hat the eifect of such large changes asubstantially autorotating rotor, the cyclic pitch n angle 0f at h 2 11 Of Subs-tan" control thereof being so coupled with the shaft tially autorotating rotor could never be compenof the rotor as to increase the ,eifective angle of fiated by chan ing he blade pitch a le B, it i the rotor control plane with the relative i t r o e con u d t at t p ph m o eonupon decrease of speed of the rotor shaft and vice 5 trolling the speed of substantially autorotating versar r cann t be a isfacto ily eels ed b the s The invention further consists in the provision of mechani m f r han ing th col ecti e blade of speed control means for a substantially auto: Dit h,- rotating rotor wherein the angle of the e'ifective Fig. 1 illustrates that adjustment of the swash rotor control plane with the relative wind may be plate with respect to the relative wind (that is,

Fig. 2 shows in plan view a simple rotor-fixed wing aircraft having a conventional fuselage B,

empennage C, and a relatively small, fixed wing D supporting conventional engines for driving the propellers E at each side of the plane of symmetry of the aircraft. In addition, the aircraft is equipped with a lifting rotor 1! having a hub I2 in which are journal-mounted the rotor blades 13, each incorporating near its root adjacent the rotor hub l2 a pitch change horn M.

The aircraft empennage C comprises a vertical or yaw-control surface 15, a horizontal stabilizer l6, and a controllable elevator surface H, which may be equipped with fixed or controllable tabs, not here shown.

Referring now to Fig. 3, the system herein provided for controlling the speed of the rotor H comprises a rotor shaft it supported for rotation in a lift bearing 19 secured to the structure of the aircraft. The upper extremity of the rotor shaft 18 terminates in the rotor hub l2, shown in Fig. 2. Mounted on the rotor shaft between the hub l2 and the lift bearing it, by means of a rotor longitudinal attitude adjusting means including a gimbel mounting or universal hinge 20, is a rotating inner swash plate 2! which is caused to rotate with the rotor H by the several links 22 connecting said swash plate 21 to the pitch-changing horns US of the respective blades l3. In Fig. 3, the swashplate 2! as represented by the reference line P and A is the angle between the plane P and the direction of relative flow V.

In mating bearing relationship with said rotating inner swash plate 21 is another part of the attitude adjusting means represented by a nonrotating outer race 23 having an aft-extending lug 24 pin-connected to the actuator push-pull rod 25 of an actuator 26 secured to the structure of the aircraft. The actuator 26, shown schematically in Fig. 3, may be of any type, either mechanical, hydraulic or electric, adapted to transmit input motion from a control member, such as the control link 21 here shown, to a linear actuator rod, such as the push-pull rod Y 25, while causing all external forces acting on such actuator rod to be absorbed in the structure of the aircraft. Such mechanical, hydraulic and electrical actuators are well-known in the art.

The control link 21, which transmits input motion to the actuator 26, is operatively pinconnected to a walking beam 23 which exerts on said control link 21 the control forces hereinafter described. The walking beam 28 constitutes a differential mechanism for transmitting control effect from a plurality of inputs to one output connected to the adjusting means 2i.

Secured to the lower end of the rotor shaft i8 is a speed governor which may be a rotating flyball governor 29 having a spring 30 whose force is adapted to resist the centrifugal force built up in the governor attendant rotation of the rotor shaft. Secured to the lower end of the governor 29 is a lug 3| pin-connected to said walking beam. Pinned to said walking beam 28 at the same point (or any other convenient point suitable for the transmission of its force) is the rod 32 of the piston 33 of a dynamic pressure pick-up 34 which operates responsive to the ram plane of the or stagnation pressure of the air stream communicated from the exterior of the aircraft by means of an external, forwardly extending pressure pick-up inlet 35 and transmitted through a pressure tube 36 into cylinder 31 of the pressure pick-up 34.

On walking beam or differential mechanism 28 at the side of the output link 21 opposite the point of connection of the governor lug 3i is operatively connected an input control pushrod 38 acting through a bellcrank 39 to form a part of the longitudinal control system 40 of the aircraft. It is evident from Fig. 3 that in the operation of said longitudinal control system, the exertion of a backward force on the control stick ll, transmitted through said longitudinal control system 46 to raise the elevator l! (which will normally cause the aircraft to nose upwardly) will tend to lower the control link 21 and produce a corresponding movement of the actuator pushpull rod 25. This will tilt the rotating swash plate 2i backward and effect a cyclic pitch change of the rotor blades [3 equivalent to tilting the plane of rotation of the rotor backward.

In flight, the automatic elements of the controls, being the governor 29 and the dynamic pressure pick-up 34, functions as follows: If the autorotating rotor is caused to decelerate, whether by increase in forward speed or other factors, such deceleration will reduce the centrifugal force of the governor 29, causing the governor lug 31 to exert a downward force on the walking beam 28 and hence a backward tilting of the rotor control plane, that is, an increase in the angle A. The rotor torque coefficient CQ will thus be augmented and the rotor speed restored. An undue increase in rotor speed will result in the governor 29 raising the walking beam 28 and lessening the angle A. With increased forward speed the ram pressure in the inlet 35 of the dynamic pressure pick-up 34 increases so as to force the piston 33 upwardly. This serves to offset the compressed force of the spring 36 of the governor 29, and there will be a consequent lessening of the angle A. The consequent lessening of the rotor lift at high forward speeds is appropriate because of the increased lift of the fixed wing.

While the governor 29 and the dynamic pressure pick-up 34 serve valuable functions at high forward speeds, it is deemed preferable that they be relieved of their functions under conditions of low speed flight. Adjacent the end of the walking beam 28 opposite the control push-rod 38, there is provided a pair of stop plates 42 drilled and mounted for sliding motion along guide rod 43. Said plates 42 are also tapped in opposite senses to receive a threaded spindle 46 whose threads run right-handed with respect to one of said stop plates Q2 and left-handed with respect to the other thereof. Secured concentrically to one end of said spindle M is a drum adapted to transmit to said spindle it the movement of a cable 46 actuated from the cabin of the aircraft by means of a cranking drum M. The walking bear 28 is equipped at its end adjacent said stop plates 42, with a bearing 43. Turning the cranking drum 4'! will cause the stop plates 42 to approach and abut said bearings 48 so that there can be no movement of the walking beam transmitted to the control link 21 independently of motion of the control stick 4|. In this manner the pilot is enabled to control the attitude of the aircraft at low speeds in a manner similar to the control of conventional helicopters.

In the range of high forward flight speeds, the stop plates 42 are moved apart so as to permit the governor 29 and associated controls to function even though the control stick 4| be held fixed.

From a consideration of the linkage to the walking beam 28. illustrated in Fig. 3, it is apparent that in fiight with the stop plates 42 withdrawn, the position of the control link 21 is a function both of the position of the control stick 4! and of the governor 29, said governor operatof the control stick 4|, cause a movement of the control link 27, the changed rotor speed resulting from change in angle caused thereby, would so act on the governor 29 as to result in substantial restoration of the original position of said control link 27. Consequently, the governor 29 tends to keep the rotor speed at a desired value independent of the position of the control stick 4|. The relative effects of the stick position and governor position on the movement of the control link 27 may be modified as desired by moving the points of connection to the walking beam 28. Further, usual types of control may be employed for the governor 29 and its spring 30 to meet the characteristics of the particular aircraft upon which the rotor speed control system is employed.

The present invention has application not merely to rotor-fixed wing aircraft but to autogyros and other forms of aircraft wherein substantially autorotating rotors may be employed. However, its advantages are most marked when employed in connection with aircraft designed to operate at high forward speeds, freeing such aircraft from the hazard of rotor deceleration at high speeds under the accelerations and disturbances of air flow likely to be encountered.

It should be understood that changes and modifications in the form, construction, arrangement and combination of the several parts of the speed control for rotor-fixed wing aircraft may be made and submitted for those herein shown and described without departing from the nature and principle of the present invention.

What I claim is:

1. An aircraft comprising a fuselage, a lifting rotor connected to said fuselage, said rotor being adjustable to vary the longitudinal attitude of the rotor plane with respect to the fuselage, a shaft connected to said rotor, rotor longitudinal attitude adjusting means connected to said rotor, actuating means for said rotor longitudinal attitude adjusting means connected to and responsive to the speed of rotation of said shaft, whereby a decrease in shaft speed will produce a decrease in forward inclination of the rotor plane, and whereby an increase in shaft speed will produce an increase in forward inclination of the rotor plane.

2. An aircraft comprising a fuselage, a lifting rotor connected to said fuselage, said rotor being adjustable to vary the longitudinal attitude of the rotor plane with respect to the fuselage, a shaft connected to said rotor, rotor longitudinal attitude adjusting means connected to said rotor, first actuating means for said rotor longitudinal attitude adjusting means connected to and responsive to the speed of rotation of said shaft, second actuating means for said rotor longitudinal attitude adjusting means comprising a ram pressure device arranged to be responsive to the forward speed of the aircraft, whereby when the forward speed of the aircraft is held fixed a decrease of the shaft speed will produce a decrease in forward inclination of the rotor plane, and whereby an increase in the shaft speed will produce an increase in forward inclination of the rotor plane.

3. An aircraft comprising a fuselage, a lifting rotor connected to said fuselage, said rotor being adjustable to vary the longitudinal attitude of the rotor plane with respect to the fuselage, a shaft shaft, manual actuating means for said rotor longitudinal attitude adjusting means, a differential mechanism having two inputs and one output, one of said inputs being connected to said automatic actuating means, the other of said being connected to said manual actuating means, said output being connected to said rotor longitudinal attitude adjusting means, whereby when said manual actuating means is held fixed a decrease in shaft speed will produce a decrease in forward inclination of the increase in shaft an increase in forward inclination of the rotor plane.

4. An aircraft comprising a fuselage, a lifting rotor connected to said fuselage, said rotor being adjustable to vary the longitudinal attitude of said manual actuating means, the other of said inputs being connected to said automatic actuating means, said output being connected to said rotor longitudinal attitude adjusting means, and stop means disposed on opposite sides of a portion of said differential mechanism to limit the movement of said portion, said stop means being adjustable to vary the differential efiect of said mechanism, whereby when said stop means are in a position to decrease in forward inclination of the rotor plane, and whereby an increase in shaft speed will produce an increase in forward inclination of the rotor plane.

References Cited in the file of this patent UNITED STATES PATENTS Number Name Date 2,421,364 Cierva May 27, 1947 2,425,651 Stalker Aug. 12, 1947 

